1. Field of the Invention
This invention relates to an apparatus and method for reducing the noise of an aircraft gas turbine engine, while at the same time improving de-icing and anti-icing performance in the area of the aircraft engine's nacelle. More particularly, this invention relates to a method and apparatus for reducing noise generated by the fan section of the aircraft engine and concurrently employing a low power electrical ice protection system to minimize ice formation on the nacelle inlet and remove any ice formed on the nacelle inlet. One or more sound-absorbing or noise abatement structures are employed at the inner surface of the nacelle inlet lip to reduce noise. The use of the low power electrical ice protection system enables the use of such structures for noise abatement in the inlet lip area while simultaneously permitting the installation and use of necessary ice protection equipment on the nacelle inlet.
2. Background Information
The desirability of reducing the noise generated by an aircraft is well-known to those skilled in the art. As discussed, for example, in U.S. Pat. No. 5,749,546, both the airplane's airframe and engine produce varying amounts of objectionable audible noise during departure and approach conditions. Aircraft engine fan noise may be suppressed at the engine nacelle inlet with a noise absorbing liner, which converts acoustic energy into heat. The liner normally consists of a porous skin supported by a honeycomb backing to provide the required separation between the porous skin and a solid backskin. This provides very effective and widely accepted noise suppression characteristics. Aircraft engines with reduced noise signatures are mandated by government authorities, and as a result, are demanded by aircraft manufacturers, airlines and local communities.
In addition, the protection of surfaces critical to the safe functioning of an aircraft during flight in icing conditions is a regulated requirement. Both the wing and the engine nacelle leading edges are similar in shape and construction, and in many aircraft designs require ice protection. With respect to the engines, ice breaking away from the engine inlet can damage various components if ingested into the engine. The formation of ice in this region also constricts the flow of air into the engine, causing losses of performance and potential engine malfunction. As a result, ice protection systems for the nacelle in general, and the engine inlet portion in particular, are needed.
In the prior art, methods for providing ice protection vary. The aircraft industry typically utilizes an ethylene glycol de-icing solution to melt the ice on the aircraft wing while the aircraft is on the ground awaiting departure. This is an expensive procedure, which may require repeated application of the glycol solution. Alternatively, a rubber tube aligned along the leading edge surface of the wing may be periodically inflated to break any ice on the wing. Most high-speed aircraft direct hot gases from the jet engine compressor onto the wing or nacelle inlet leading edges to melt the accreted ice. This option is not available for aircraft that do not have jet engines, or aircraft limited in the amount of hot air that can be bled off the engine.
U.S. Pat. No. 6,027,075 describes a system for modifying the adhesion strength of ice adhered to an object. The system includes an electrode that is electrically insulated from the object, and a direct current source coupled to the object and the electrode to generate a direct current bias to an interface between the ice and the object. The direct current bias has a voltage that modifies the ice adhesion strength selectively. The electrode may include a grid electrode shaped to conform to a surface of the object. Each point of the grid electrode is in electrical contact with the source.
One example of an aircraft engine nacelle ice protection system that uses hot air bleed from the compressor stage of the turbofan engine and routes the air to the leading edge surface is the ROHRSWIRL™ system described, for example, in U.S. Pat. No. RE 36,215. In this system, tubes or ducts direct hot air from the compressor section of the engine into the D-duct, which is the cavity formed by the nacelle leading edge skin and the inlet forward bulkhead. Hot air is circulated around the perimeter of the structure and exits at a lower vent. A control valve within the system provides the means for commanding the delivery of bleed air when the ice protection system is required.
However, the use of engine bleed air for ice protection can negatively impact engine performance, and dictates that additional engine nacelle design requirements and conditions be considered. Utilizing compressed air for ice protection is an energy loss and results in propulsion system inefficiency. Modern turbofan engines increase thrust levels by increasing the by-pass ratio, i.e. the amount of fan flow divided by the amount of core flow. To do this, fan diameters have been increased, thus resulting in a larger D-duct volume and a greater surface area requiring ice protection capability. At the same time, the availability of core compressor air for de-icing and anti-icing has been decreased. The resulting requirement of hot air ice protection systems is a significant performance penalty incurred by the engine on take-off.
Furthermore, because the ducting that carries the hot bleed air passes through various engine compartments on its way to the D-duct, consideration must be given to duct failure modes, which requires that pressure relief doors be added. Additionally the surrounding structure must be able to operate at high temperatures. Consideration must also be given to the use of materials having different coefficients of thermal expansion. Material selection for the inlet lip assembly must take into account the conditions of a high temperature and pressure environment, plus the need for structural integrity in the event of a bird strike or other foreign object damage. The bulkhead is typically made from nickel-based INCONEL® alloys or titanium alloys, with the lipskin being made from a high temperature aluminum alloy. The lip assembly interfaces with the outer barrel and inner acoustic bond panel, which can be manufactured from either a composite or aluminum material. Where these materials join together, care must be taken to account for the differing coefficients of thermal expansion, the conduction of heat into lower temperature resistant structures, as well as galvanic corrosion requirements, as are well understood by those skilled in the art.
In view of the foregoing, it is necessary to both suppress noise at the engine nacelle inlet and provide sufficient ice protection capability for the leading edges of the nacelle. However, the portion of the engine nacelle inlet lip available for acoustic treatment is currently limited to the inner barrel structure aft of the inlet lip, due mainly to the incompatibility of the prior art hot air de-icing systems with the relatively low temperature capability of existing, adhesively bonded honeycomb noise abatement structures. There is a lack of current technology that provides a satisfactory method for extending such engine noise reduction structures into the nacelle lip while concurrently satisfying the temperature constraints of the noise abatement treatment. Thus, there is a need in the art for a method and apparatus that satisfies the requisite nacelle inlet ice protection requirements while at the same time permitting the use of noise abatement structures within the nacelle inlet lip to achieve the noise reduction required of modern gas turbine powered aircraft.